The disclosed embodiments generally pertain to fan sections of an aircraft engine. More particularly, but not by way of limitation, the present embodiments relate to hybrid aircraft components formed of two or more parts of differing composition.
A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet or intake is at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by a compressor, a combustion chamber, a turbine, and a nozzle at the aft end of the engine. It will be readily apparent from those skilled in the art that additional components may also be included in the engine, such as, for example, low-pressure and high-pressure compressors, and high-pressure and low-pressure turbines. This, however, is not an exhaustive list. An engine also typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The internal shaft is connected to both the turbine and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades.
In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. In a two stage turbine, a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy. The second stage turbine blades and rotor disk are mechanically coupled to a low pressure or booster compressor for driving the booster compressor and additionally an inlet fan.
It is always desirable to reduce the weight of a gas turbine engine and its related components, especially for those utilized in the aviation industry. Within the fan section of the engine, current support structures mounted aft of the spinner or cone are generally manufactured of titanium. Due to the use of titanium, such support structure is a relatively heavy material and also expensive. It would be desirable to replace this titanium material with a less expensive and lighter weight material without affecting the hoop load characteristics of the support ring. However, such spinner support ring cannot be manufactured from a single material of lower density, such as aluminum, due to material limitations or capabilities. Thus, while it would be desirable to reduce the cost and weight of the support ring, it is also necessary to maintain the load carrying capability of the titanium part being replaced.
As may be seen by the foregoing, it would be desirable to overcome these and other deficiencies with gas turbine engine components.